This application is based on and claims the priority under 35 U.S.C. xc2xa7119 of German Patent Application 101 63 848.5, filed on Dec. 22, 2001, the entire disclosure of which is incorporated herein by reference.
The invention relates to a metal structural component for an aircraft, including a plurality of stiffening profile members arranged on and integrally connected with a metal skin sheet.
The majority of aircraft structures, such as the fuselage and the wings for example, are fabricated of metal or metallic materials. In order to satisfy primarily the stiffness and strength requirements, the prevailing conventional construction technique for fabricating such aircraft structures is to provide stiffening members on a metal skin sheet. An especially meaningful example thereof is the fuselage, of which the outer fuselage skin is reinforced or strengthened by stringers extending in the longitudinal direction and by frames extending in the circumferential direction. Typically, riveting or adhesive bonding is used as the joining method for joining the stringers and/or frames onto the skin sheets. Such a construction method is generally referred to as a differential construction method, because it involves a non-integral connection between the stiffening members and the skin sheets as separate differentiated components.
Furthermore, it is also known to fabricate integral structural components with a direct integral connection between the stiffening members and the skin. This integral construction method may, for example, involve simultaneously extruding the skin and the stiffening members as a single integral component, or the use of welding methods for integrally joining the stiffening members onto the skin in order to form a resulting integral structural component. For example, it is known to use laser beam welding for fabricating a structural component including stringers welded onto a skin, for example as disclosed in German Patent DE 196 39 667 C1 and corresponding U.S. Pat. No. 5,841,098.
Structural components of the above described type are typically used as fuselage shells in the construction of an aircraft. During the subsequent operation of the aircraft, both static as well as cyclical mechanical loads arise in the fuselage, which can lead to the formation of cracks or rips in the reinforced, i.e. strengthened, metal skin sheet. Once such a crack or rip forms in the skin sheet, it tends to propagate along the skin sheet. In consideration of such potential damage, both the requirements of the crack propagation characteristics as well as the residual or remaining strength characteristic of the structural component are taken into account during the design of the aircraft, and especially the pertinent structural components thereof. The required damage tolerance characteristic of such a structural component is dependent on how a skin crack will behave once it forms in and propagates along the skin sheet and then meets perpendicularly onto a stiffening member.
In the case of a differential construction method, in which the stiffening members are riveted or adhesively bonded onto the skin sheets, a skin crack will typically continue propagating and pass directly under the stiffening member, without propagating into the stiffening member. Thus, the stiffening member thereby remains undamaged, while bridging across and holding together the crack in the underlying skin sheet, so that as a positive effect, the further propagation of the crack is stopped or hindered. In other words, in such a differential manner of construction, the riveted or adhesively bonded stiffening members act as crack or rip stoppers or at least crack propagation retarding elements. Thereby, the further crack propagation in the fuselage shell is stopped, hindered, or retarded, because the front tip of the propagating crack will be held together, i.e. the two portions of the skin sheet on opposite sides of the crack will be held together by the riveted or adhesively bonded stiffening member, at least for a certain number of load cycles.
On the other hand, in the case of an integral manner of construction, in which the stiffening members are integrally connected with the skin sheets, a crack propagating in the skin sheet, once it reaches a stiffening member, will also propagate into the stiffening member. In other words, the crack front or tip, once it reaches the stiffening member, will be divided into a skin crack running in the skin sheet and a stiffening member crack running in the base or pedestal portion of the stiffening member. Both of these cracks will then continue to propagate in the two respective component portions independently of each other.
Such an integral stiffening member that has been damaged by a partial crack therein thus has both a reduced strength as well as a reduced stiffness in comparison to a stiffening profile member that has been connected to the skin in a differential non-integral manner. As a result of the cracking and thus weakening of the integral stiffening member, it no longer provides an effective holding-together of the skin sheet portions on opposite sides of the crack, so that the further propagation of the crack is not adequately prevented or retarded. This cracking process leads to a reduced residual or remaining strength of the structural component after a crack has been formed therein, and to an overall disadvantageous crack propagation characteristic. In critical areas of the fuselage shell in which the residual or remaining strength is determinative as a design criterium, the fuselage shells must be thickened to provide greater strength and stiffness, and especially an adequate residual or remaining strength even after the crack has propagated therein. This in turn leads to an unacceptable weight increase, especially in the side and overhead shell areas of the fuselage.
German Patent DE 199 24 909 C1 discloses an approach to solving or addressing the above described problem in an integral structural component. Particularly, this reference discloses a thickening in the base or pedestal portion of the stiffening member of the structural component, for retarding or deflecting the further propagation of a crack. However, since the thickening of the pedestal portion of the stiffening member is integral with the stiffening member, i.e. the thickening is an integral common portion (i.e. integrally joined on the atomic level) of the stiffening member, there basically still exists a danger that a crack can propagate unchecked through the entire cross-section of the thickening member. Moreover, the external shape of such a stiffening member differs from the previously conventional stiffening member shapes, due to the extra protruding thickened area of the pedestal portion. Therefore, tools, jigs and other devices used during the fabrication of an aircraft fuselage must in many cases be altered or adapted to the changed configuration of the stiffening members, which leads to additional tooling and set-up costs.
In view of the above, it is an object of the present invention to provide a metallic structural component including stiffening members integrally connected with a skin sheet, wherein a large surface propagation of a crack is prevented or hindered, and therewith an improved crack propagation characteristic and an increased residual strength after formation of a crack are achieved. Further objects of the invention are to provide such a structural component that meets the requirements of a lightweight structure as needed in the field of aircraft construction, and to provide a stiffening profile member that has an improved crack propagation resistance yet has an external cross-sectional shape that is essentially the same as any desired conventional stiffening profile member without the special inventive features. The invention further aims to avoid or overcome the disadvantages of the prior art, and to achieve additional advantages, as apparent from the present specification.
The above objects have been achieved according to the invention in a metallic structural component for an aircraft including a metal skin sheet and a plurality of stiffening profile members integrally connected onto the skin sheet, wherein each stiffening profile member comprises a first part that is integrally connected to the skin sheet and a second part that is non-integral with the first part, so as to form an internal boundary surface between the first part and the second part. The internal boundary surface acts to prevent or hinder a further crack propagation across this internal boundary surface. The overall cross-sectional shape of the stiffening profile member is made up of the two parts thereof, whereby this overall cross-sectional profile shape may have essentially the same shape as any conventionally known stiffening profile member. In other words, the provision of the separate first and second parts of the stiffening profile member does not significantly influence or change the overall cross-sectional profile shape of the stiffening profile member.
With the inventive structure of the stiffening profile member, it is especially advantageous that the post-crack residual strength of a metallic structural component having an integral construction is increased, so that such a structural component is suitable for use as a fuselage shell of an aircraft even in areas that are primarily loaded under tension and that are therefore critical in view of the fatigue characteristics of the component. Thus, it now becomes possible to use welded or otherwise integral fuselage shells for the entire fuselage of the aircraft, which advantageously reduces the effort and expense of the construction of the aircraft, due to the advantages in the fabrication. Moreover, the crack stopping or retarding measures are incorporated within the usual previously-existing cross-sectional profile shape of the stiffening profile member, and do not have a substantial effect on the stiffening member geometry (for example not requiring an external thickening or reinforcement of the previously existing stiffening members). It is thus possible to continue to use all of the previously existing manufacturing equipment and methods, without any special changes.
The inventive structure eliminates or substantially overcomes the disadvantages of the crack propagation characteristic of welded shell components, i.e. integral shell components. In the event of the occurrence of a primary crack in the fuselage skin, the simultaneously formed crack in the stiffening profile member will be stopped, deflected, or adequately hindered or retarded from further propagation in the stiffening profile member upon reaching the internally incorporated boundary surfaces between the first part and the second part of the respective stiffening profile member.